The subject matter disclosed herein relates to gas turbine engines and, more particularly, to temperature management therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gas that flows downstream through one or more turbine stages. A turbine stage includes a stationary turbine nozzle having stator vanes that guide the combustion gas through a downstream row of turbine rotor blades. The blades extend radially outwardly from a supporting disk that is powered by extracting energy from the gas.
A first stage turbine nozzle receives hot combustion gas from the combustor that is directed to the first stage turbine rotor blades for extraction of energy therefrom. A second stage turbine nozzle may be disposed downstream from the first stage turbine rotor blades, and is followed by a row of second stage turbine rotor blades that extract additional energy from the combustion gas. Additional stages of turbine nozzles and turbine rotor blades may be disposed downstream from the second stage turbine rotor blades.
As energy is extracted from the combustion gas, the temperature of the gas is correspondingly reduced. However, since the gas temperature is relatively high, the turbine stages are typically cooled by diverting air from the compressor through the hollow vane and blade airfoils as well as sidewalls and shrouds. Since the cooling air is diverted from use by the combustor, the amount of extracted cooling air has a direct influence on the overall efficiency of the engine. It is therefore desired to improve the efficiency with which the cooling air is utilized to improve the overall efficiency of the turbine engine.
The quantity of cooling air required is dependant on the temperature of the combustion gas. Since combustion gas temperature directly affects the gas turbine component capability of meeting operating life requirements, the cooling air requirement for the turbine stages must be effective for withstanding high temperature operation of the engine.
The combustion gas temperature varies temporally over the operating or running condition of the engine and also varies circumferentially based on the location at which the gas is discharged from the outlet of the combustor. Large circumferential temperature variations are particularly present in can-annular combustion systems, where outlets of multiple combustion cans form the annular combustion outlet. The combustion gas temperature peaks in the center of each can outlet while the temperature at the sides of the can outlet is lower due to combustion aft-frame leakage. This spatial temperature variation is typically represented by combustor pattern and profile factors that are conventionally known.
Accordingly, the stationary components of each turbine stage are specifically designed for withstanding the peak combustion gas temperature. Since the segments in each row of vane airfoils, vane sidewalls and shrouds are often similar to each other, the cooling configurations may also be similar. As a result, the cooling configurations are effective for providing suitable cooling at the peak combustion gas temperatures experienced by the individual stages. Each vane airfoil, vane sidewall and shroud is cooled based on the peak temperature on the combustor pattern profile. This results in excess cooling for segments located downstream of lower temperature regions of the combustor outlet. Excess cooling translates directly to lower than desired turbine efficiency.
It is therefore desired to provide a gas turbine engine having improved cooling of gas turbine stationary components.